The primary structural elements of many aircraft, typically larger aircraft such as large passenger jets, are often made from metal. Fuselage shells and sections for such aircraft, for example, are typically manufactured from high-strength aluminum alloys or similar metals. Recently, however, many aircraft manufacturers have begun using composite materials, such as fiber-reinforced resin materials, for the relatively high strength-to-weight ratios of such materials. Suitable composite materials usually include glass, carbon, or synthetic (e.g. polyamide, polyester, etc.) fibers in a matrix of epoxy or another type of resin.
One example method of manufacturing aircraft fuselages with composite materials involves wrapping fibers or fiber tapes around a rotating mandrel, generally with an automated instrument or system. The mandrel provides the basic shape of a longitudinal fuselage section. The fibers or tapes may be pre-impregnated with an epoxy, or passed through a resin bath just before the material contacts the rotating mandrel, and are applied in multiple plies to form a skin of a fuselage section. In some techniques, the mandrel can remain in place and become part of the wound component, or it can be removed. The skin may be covered with additional layers, such as a layer of honeycomb core, to which additional plies of composite materials may be applied in a composite “sandwich” structure.
One way in which composite fuselage sections formed in this manner are joined together involves the use of one or more splice plates that are fastened in place over a circumferential splice joint between adjacent fuselage sections. In general, a splice plate is held in place while attachment holes are precisely drilled through the splice plate and underlying composite structure. Each hole is usually probed for size quality, such as through the use of a machine that may also record statistical process control data on each hole. Fasteners are then applied according to precisely measured torques, securing the plate in place.
For larger aircraft, such as widebody aircraft, a compound splice plate system may be used. Typically, a system of structural beams such as longitudinal stiffeners and lateral frames are attached to the fuselage sections for reinforcement. A compound splice plate system may involve splice plates having longitudinal extensions positioned along the splice plate to extend to the regions on either side of the splice plate between adjacent stiffeners, and/or separate longitudinal fittings that are fastened cross-wise to the splice plate to extend to either side and in between adjacent stiffeners. The extensions and/or fittings are fastened to the composite structure in a similar manner as the splice plate. In some cases, the extensions or fittings are provided with orthogonal flanges, such as for additional rigidity, which may themselves be attached to stiffeners and/or frames.
Methods of splice plate joining of composite fuselage sections thus can involve labor-intensive assembly procedures and extensive tooling fixtures, especially in widebody aircraft. For example, the inner diameter of the fuselage in widebody aircraft is typically around 200 inches (around 5 meters). The number of fasteners required for joining adjacent fuselage sections of this size can exceed 5,000. Even with automation, the time required to drill the requisite number of holes and properly apply the fasteners can be over 400 man-hours per section. Moreover, widebody aircraft fuselages are assembled from several fuselage sections—for example, the Boeing 787 Dreamliner is assembled from five fuselage sections. As such, the aforementioned joining technique for the fuselage sections for this size and type of aircraft can represent a demand for a fastener count exceeding 20,000, and over 2,000 man-hours.